Abstract: A fan blade platform assembly includes a pressure side platform half that includes a first attachment wall, a first gas path wall, and a first connecting wall attached to both the first attachment wall and the first gas path wall. A suction side platform half includes a second attachment wall, a second gas path wall, and a second connecting wall attached to both the second attachment wall and the second gas path wall. A seal engages the pressure side platform half and the suction side platform half.

Abstract: A regenerative blower-compressor includes an impeller mounted to a drive shaft within a housing including a channel extending from an inlet adjacent to a low fluid-pressure region of the channel to an outlet adjacent to a high fluid-pressure region of the channel, the impeller extends radially outward through an annular volume within the housing from the drive shaft to blades in the channel and is configured to rotate for rotating the blades through the channel for forcing fluid through the channel from the inlet to the outlet in response to rotation of the drive shaft, the drive shaft extends from the impeller within the annular volume to into a shaft chamber within the housing configured to receive fluid from the high fluid-pressure region of the channel, and a port configured to vent fluid directly from the shaft chamber to into the low fluid-pressure region of the channel.

Abstract: The invention relates to a straightening assembly (28) comprising two radially inner and outer coaxial shells (34) between which extend vanes (36) made of composite material, fixed to a first end portion on the radially inner shell and to a second end portion on the radially outer shell, characterized in that: for each vane (36), in a plane perpendicular to the axis of the radially inner and outer shells (34), a straight line passing through a junction between said first and second end portions and the useful part forms an angle ? with a radius of the radially inner shell, passing through the junction between said first end portion and the useful part of said vane (36), such that 0°<??30°; the radially inner shell has a diameter ranging from 1,000 mm to 1,600 mm; and in that the radially outer shell has a diameter ranging from 2,000 to 2,800 mm; the number of vanes (36) ranges from twenty-five to forty-five.

Abstract: Disclosed herein are a grooved turbine blade, and a turbine and a gas turbine including the same. According to the disclosure, since a groove part is formed on a root member to disperse torsional stress, it is possible to improve durability of a rotor disk to which the turbine blade is assembled.

Abstract: A turbine comprises a casing, an outer metal shroud, an inner metal shroud and an annular distributor having a plurality of CMC ring sectors, each sector comprising a mast, an inner platform, an outer platform and at least one blade having a hollow profile that defines an inner housing, the inner and outer platforms each having an opening communicating with said inner housing, and the mast passing through said openings and the inner housing and being secured to said casing and connected to said annular sector. Each blade comprises at least one first radial shoulder projecting axially towards the inside of the blade, and each mast comprises at least one second shoulder projecting axially towards the outside of the mast configured to radially cooperate with a first shoulder and radially press the blade against the mast.

Abstract: A retention ring for a gas turbine engine according to an example of the present disclosure includes, among other things, a main body extending in a circumferential direction about an axis to establish a continuous hoop having a first diameter and a second diameter. The main body includes first and second circumferential faces along opposite sides of the main body. The first circumferential face is dimensioned to abut a gas turbine engine component. The main body includes at least one removal feature dimensioned to sever in response to engagement with a cutting tool. A method of assembly is also disclosed.

Abstract: A turbomachine defining a flowpath therethrough at which a fluid is compressed is provided, in which the turbomachine includes a first compressor in serial flow arrangement upstream of a second compressor. The second compressor includes a port at the flowpath and is configured to receive at least a portion of the fluid from the flowpath from the second compressor. The first compressor includes a vane positioned at the flowpath and the vane includes an opening at the flowpath configured to egress the portion of the fluid from the port into the flowpath.

Abstract: A turbine nozzle assembly includes at least one airfoil and an outer endwall coupled to a radial outer end of the at least one airfoil. A mounting rail is coupled to the outer endwall and extends at least partially radially outward from the outer endwall and at least partially circumferentially along the outer endwall. A stress relief structure is defined in the mounting rail. The stress relief structure includes an end opening defined in a radial outer surface of the mounting rail, a slot defined through the rail thickness of the mounting rail and coupled to the end opening, and an oblong opening defined through the rail thickness of the mounting rail and coupled to a radial inner end of the slot. The oblong opening is arranged asymmetrically in a circumferential direction relative to the slot to relieve stress where prevalent in the mounting rail.

Abstract: A turbomachine diaphragm including a sealing section having a first end portion that extends to a second end portion through an intermediate portion; and at least one rail member including a first end section that extends from the first end portion of the sealing section to a second end section through an intermediate section having an inner surface section and an outer surface section, the second end section including multiple weld passes disposed on opposed sides of the second end section for mitigation of thermal tensions on the diaphragm, the multiple weld passes forming a cladding welded to the diaphragm, wherein the cladding includes a stainless austenitic steel.

Abstract: A rotor vane for an aircraft turbine engine includes a blade extending between an inner platform and an outer platform which carries at least one projecting lip. The blade has a lower surface and an upper surface, and the outer platform includes, on the side of the lower and upper surfaces, lateral edges configured to cooperate in a form-fitting manner with complementary lateral edges of adjacent vanes. Each of the lateral edges has an anti-wear coating, and one of the lateral edges forms a hollow tip (P) with a bowl configured to receive the coating and further including a first concave cylindrical surface portion, the geometric dimensions of which are optimized to limit the risk of cracks appearing when the coating is applied.

Abstract: A method for determining angular position of a generator shaft of a wind turbine is presented. The method includes receiving a position signal, from an encoder position sensor of the wind turbine, indicative of an angular position of the generator shaft. The method includes determining a compensation signal to compensate for a disturbance signal in the received position signal indicative of an imperfection associated with the encoder position sensor. The method includes modifying the position signal by applying the determined compensation signal to the received position signal to determine angular position.

Abstract: A method of verifying operation of an aircraft in a pre-defined flight mode includes operating the aircraft in a first flight mode, commanding the aircraft to transition to a second flight mode, evaluating a plurality of motor performance parameters, and based on values of the plurality of motor performance parameters, determining whether the aircraft has successfully transitioned to the second flight mode.

Abstract: A runner for a hydraulic machine comprising a band, a crown, a plurality of blades extending between the crown and the band, wherein the runner comprises a plurality of runner segments which together define the runner, each runner segment comprising a band portion, a crown portion and a blade, which portions are integrally formed with one another, each runner segment being attachable to another segment at a band joining edge and a crown joining edge, wherein the band joining edge and the crown joining edge are each spaced apart from the blade of the segment.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction. An inner core engine has an inner core engine housing surrounding a compressor section, including a low pressure compressor. A rigid connection between a fan case and the inner core engine includes A-frames rigidly connected at a connection point to the fan case. Fan exit guide vanes rigidly connect to the fan case, and to the inner core engine. A fan intermediate case is positioned forward of a first rotor stage in the low pressure compressor. A rigid structure is connected to the inner core engine and to the fan exit guide vanes. The rigid structure defines a structure moment stiffness. The fan intermediate case defines an intermediate case moment stiffness. A ratio of the structure moment stiffness to the intermediate case moment stiffness is between 5 and 15.

Abstract: A vane multiplet includes a plurality of airfoils that each have a flared end, a common platform piece, a plurality of seal, and a plurality of retainers. The common platform has airfoil sockets that each define an airfoil opening that is circumscribed by a groove. The flared ends of the airfoils are seated in the airfoil sockets such that the grooves and the airfoils together form seal channels that have an open side. The seals are disposed in the seal channels. The retainers have airfoil-shaped profiles and are disposed in the airfoil sockets to bound the open sides of the seal channels and retain the seal in the seal channel.

Abstract: A variable guide vane (VGV) assembly, has: vanes distributed about a central axis, each of the vanes having an airfoil extending between first and second stems, the vanes rotatable about vane axes extending between the first and second stems; and a bushing ring supporting the first stems of the vanes, the bushing ring defining circumferentially distributed pockets receiving the first stems of the vanes, the bushing ring having a first ring body extending circumferentially around the central axis, a second ring body extending circumferentially around the central axis and removably secured to the first ring body, the second ring body defining at least one tab projecting axially to axially overlap at least a portion of the first ring body and defining a contact plane therebetween, and at least one locking member extending radially through the contact plane to removably secure the first ring body and the second ring body together.

Abstract: A steam turbine has a diffuser that is configured to guide steam to an outside of a casing. The diffuser has an outer guide that gradually expands to an outer side in a radial direction and an inner guide that is disposed at intervals to an inner side in the radial direction with respect to the outer guide. The inner guide has an inner curved diameter-expanded portion that gradually expands to the outer side in the radial direction while being curved from the first side to the second side in the axial direction. The outer guide has a first diameter-expanded portion that gradually expands to the outer side in the radial direction with a first radius of curvature, and a second diameter-expanded portion that gradually expands to the outer side in the radial direction with a second radius of curvature larger than the first radius of curvature.

Abstract: An aircraft engine has: a first component and a second component coaxially mounted about a central axis; a flow passage extending within an annular gap defined radially between the first component and the second component, the flow passage fluidly connecting a first zone to a second zone; a seal disposed in the flow passage between the first zone and the second zone, the seal extending from a seal base secured to the first component to a seal end radially spaced apart from the seal base; a deflector located downstream of the seal relative to a first flow flowing from the first zone to the second zone through the seal, the deflector extending from a deflector base secured to the second component to a deflector end radially spaced apart from the deflector base; and a radial gap defined radially between the seal end and the deflector end.

Abstract: A turbine rotor in an embodiment is configured by joining a rotor component member and a rotor component member together by bolt fastening with an abutting end surface of the rotor component member and an abutting end surface of the rotor component member abutting on each other. The turbine rotor includes: a cylindrical recessed portion that is formed at the abutting end surface and is recessed in an axial direction; an axial passage bored from a bottom surface of the cylindrical recessed portion in the axial direction; an introduction passage introducing the cooling medium into the axial passage; a discharge passage discharging the cooling medium from the axial passage; and a sealing member that is arranged in the cylindrical recessed portion and seals one end of the axial passage.

Abstract: A hydrostatic advanced low leakage seal configured to be disposed between relatively rotatable components. The seal includes a base. The seal also includes a shoe extending circumferentially. The seal further includes a radially outer beam operatively coupling the shoe to the base. The seal yet further includes a radially inner beam operatively coupling the shoe to the base, wherein one of the radially inner beam and the radially outer beam is oriented to be angled relative to the other of the radially inner and outer beam.